Gas turbine engine compressor arrangement

ABSTRACT

A turbofan gas turbine engine includes, among other things, a fan section, a core engine, a bypass passage, and a bypass ratio defined as the volume of air passing into the bypass passage compared to the volume of air passing into the core engine, the bypass ratio being greater than or equal to 10. A gear arrangement drives the fan section. A compressor section includes a low pressure compressor section and a high pressure compressor section. A turbine section drives the gear arrangement. An overall pressure ratio is provided by the combination of a pressure ratio across the low pressure compressor section and a pressure ratio across the high pressure compressor section, and greater than 40. The pressure ratio across the high pressure compressor section is between 7 and 15, and the pressure ratio across the low pressure compressor section is between 4 and 8.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a continuation of U.S. application Ser. No.15/184,253, filed Jun. 16, 2016, which is a continuation of U.S.application Ser. No. 14/179,640, filed Feb. 13, 2014, which is acontinuation-in-part of U.S. application Ser. No. 13/869,057, filed Apr.24, 2013 (now U.S. Pat. No. 9,121,367), which is a continuation of U.S.application Ser. No. 13/590,273, filed Aug. 21, 2012 (now U.S. Pat. No.8,449,247), which is a continuation of U.S. application Ser. No.13/418,457, filed Mar. 13, 2012 (now U.S. Pat. No. 8,277,174), whichclaims priority to U.S. Provisional Application 61/604,646, filed Feb.29, 2012, and is a continuation in-part of U.S. patent application Ser.No. 13/337,354, filed on Dec. 27, 2011 (now U.S. Pat. No. 8,337,147),and entitled “Gas Turbine Engine Compressor Arrangement,” which was acontinuation-in-part of U.S. patent application Ser. No. 13/294,492filed on Nov. 11, 2011, and entitled “Gas Turbine Engine Compressor CaseMounting Arrangement,” which was a continuation of U.S. patentapplication Ser. No. 11/858,988 filed on Sep. 21, 2007 (now U.S. Pat.No. 8,075,261), and entitled “Gas Turbine Engine Compressor CaseMounting Arrangement.”

BACKGROUND

The present invention relates generally to a gas turbine engine.

Gas turbine engines are known, and typically include a compressor forcompressing air and delivering it downstream into a combustion section.A fan may move air to the compressor. The compressed air is mixed withfuel and combusted in the combustion section. The products of thiscombustion are then delivered downstream over turbine rotors, which aredriven to rotate and provide power to the engine.

The compressor includes rotors moving within a compressor case tocompress air. Maintaining close tolerances between the rotors and theinterior of the compressor case facilitates air compression.

Gas turbine engines may include an inlet case for guiding air into acompressor case. The inlet case is mounted adjacent the fan section.Movement of the fan section, such as during in-flight maneuvers, maymove the inlet case. Some prior gas turbine engine designs support afront portion of the compressor with the inlet case while anintermediate case structure supports a rear portion of the compressor.In such an arrangement, movement of the fan section may cause at leastthe front portion of the compressor to move relative to other portionsof the compressor.

Disadvantageously, relative movement between portions of the compressormay vary rotor tip and other clearances within the compressor, which candecrease the compression efficiency. Further, supporting the compressorwith the inlet case may complicate access to some plumbing connectionsnear the inlet case.

It would be desirable to reduce relative movement between portions ofthe compressor and to simplify accessing plumbing connection in a gasturbine engine.

Traditionally, a fan and low pressure compressor have been driven in oneof two manners. First, one type of known gas turbine engine utilizesthree turbine sections, with one driving a high pressure compressor, asecond turbine rotor driving the low pressure compressor, and a thirdturbine rotor driving the a fan. Another typical arrangement utilizes alow pressure turbine section to drive both the low pressure compressorand the fan.

Recently it has been proposed to incorporate a gear reduction to drivethe fan such that a low pressure turbine can drive both the low pressurecompressor and the fan, but at different speeds.

SUMMARY

A gas turbine engine, according to an exemplary aspect of the presentdisclosure includes, among other things, a fan section including a fanand a gear arrangement configured to drive the fan section. A compressorsection includes both a first compressor and a second compressor. Aturbine section is configured to drive the compressor section and thegear arrangement. An overall pressure ratio is provided by thecombination of a pressure ratio across the first compressor and apressure ratio across the second compressor and is greater than or equalto about 35. The pressure ratio across the first compressor is greaterthan or equal to about 7. A pressure ratio across the fan section isless than or equal to about 1.50. The fan is configured to deliver aportion of air into the compressor section, and a portion of air into abypass duct.

In a further non-limiting embodiment of the foregoing gas turbineengine, the first compressor is upstream of the second compressor.

In a further non-limiting embodiment of either of the foregoing gasturbine engines, the first compressor is downstream of the secondcompressor.

In a further non-limiting embodiment of any of the foregoing gas turbineengines, the pressure ratio across the fan section is less than or equalto about 1.45.

In a further non-limiting embodiment of any of the foregoing gas turbineengines, the overall pressure ratio is above or equal to about 50.

In a further non-limiting embodiment of any of the foregoing gas turbineengines, the geared arrangement defines a gear reduction ratio greaterthan or equal to about 2.3.

In a further non-limiting embodiment of any of the foregoing gas turbineengines, a bypass ratio, which is defined as a volume of air passing tothe bypass duct compared to a volume of air passing into the compressorsection, is greater than or equal to about 8.

In a further non-limiting embodiment of any of the foregoing gas turbineengines, the turbine section includes a fan drive turbine configured todrive the fan section, a pressure ratio across the fan drive turbinebeing greater than or equal to about 5.

In a further non-limiting embodiment of any of the foregoing gas turbineengines, the fan section includes a plurality of fan blades and a fanblade tip speed of each of the fan blades is less than about 1150ft/second.

A gas turbine engine, according to an exemplary aspect of the presentdisclosure includes, among other things, a fan section including a fan,and a gear arrangement configured to drive the fan section. A compressorsection includes both a first compressor and a second compressor. Aturbine section is configured to drive the compressor section and thegear arrangement. An overall pressure ratio is provided by thecombination of a pressure ratio across the first compressor and apressure ratio across the second compressor and is greater than or equalto about 35. The pressure ratio across the first compressor is less thanor equal to about 8. A pressure ratio across the fan section is lessthan or equal to about 1.50. The fan is configured to deliver a portionof air into the compressor section, and a portion of air into a bypassduct.

In a further non-limiting embodiment of the foregoing gas turbineengine, the first compressor is upstream of the second compressor.

In a further non-limiting embodiment of either of the foregoing gasturbine engines, the first compressor is downstream of the secondcompressor.

In a further non-limiting embodiment of any of the foregoing gas turbineengines, the pressure ratio across the fan section is less than or equalto about 1.45.

In a further non-limiting embodiment of any of the foregoing gas turbineengines, the overall pressure ratio is above or equal to about 50.

In a further non-limiting embodiment of any of the foregoing gas turbineengines, the pressure ratio across the second compressor is greater thanor equal to about 7.

In a further non-limiting embodiment of any of the foregoing gas turbineengines, the pressure ratio across the first compressor is between about3 and about 8, and the pressure ratio across the second compressor isbetween about 7 and about 15.

In a further non-limiting embodiment of any of the foregoing gas turbineengines, the geared arrangement defines a gear reduction ratio greaterthan or equal to about 2.3.

In a further non-limiting embodiment of any of the foregoing gas turbineengines, a bypass ratio, which is defined as a volume of air passing tothe bypass duct compared to a volume of air passing into the compressorsection, being greater than or equal to about 8.

In a further non-limiting embodiment of any of the foregoing gas turbineengines, the turbine section includes a fan drive turbine configured todrive the fan section, a pressure ratio across the fan drive turbinebeing greater than or equal to about 5.

In a further non-limiting embodiment of any of the foregoing gas turbineengines, the fan section includes a plurality of fan blades and a fanblade tip speed of each of the fan blades is less than about 1150ft/second.

A gas turbine engine, according to an exemplary aspect of the presentdisclosure includes, among other things, a fan section including a fan,and a gear arrangement configured to drive the fan section. A compressorsection includes both a first compressor and a second compressor. Aturbine section is configured to drive the compressor section and thegear arrangement. An overall pressure ratio is provided by thecombination of a pressure ratio across the first compressor and apressure ratio across the second compressor, the pressure ratio acrossthe first compressor being less than about 8, and the pressure ratioacross the second compressor being greater than or equal to about 7. Apressure ratio across the fan section is less than or equal to about1.50. The fan is configured to deliver a portion of air into thecompressor section, and a portion of air into a bypass duct.

In a further non-limiting embodiment of the foregoing gas turbineengine, the first compressor is upstream of the second compressor.

In a further non-limiting embodiment of the foregoing gas turbineengine, the first compressor is downstream of the second compressor.

In a further non-limiting embodiment of the foregoing gas turbineengine, the pressure ratio across the fan section is less than or equalto about 1.45.

In a further non-limiting embodiment of the foregoing gas turbineengine, the overall pressure ratio is greater than or equal to about 35.

In a further non-limiting embodiment of the foregoing gas turbineengine, the geared arrangement defines a gear reduction ratio greaterthan or equal to about 2.3.

In a further non-limiting embodiment of the foregoing gas turbineengine, a bypass ratio, which is defined as a volume of air passing tothe bypass duct compared to a volume of air passing into the compressorsection, being greater than or equal to about 8.

In a further non-limiting embodiment of the foregoing gas turbineengine, the turbine section includes a fan drive turbine configured todrive the fan section, a pressure ratio across the fan drive turbinebeing greater than or equal to about 5.

An arrangement for a gas turbine engine, according to an exemplaryaspect of the present disclosure includes, among other things, a fansection having a central axis, a compressor case for housing acompressor, and an inlet case for guiding air to the compressor, thecompressor case positioned axially further from the fan section than theinlet case. A support member extends between the fan section and thecompressor case wherein the support member restricts movement of thecompressor case relative to the inlet case. The compressor case includesa front compressor case portion and a rear compressor case portion, therear compressor case portion being axially further from the inlet casethan the front compressor case portion. The support member extendsbetween the fan section and the front compressor case portion, and theinlet case is removable from the gas turbofan engine separately from thecompressor case. The compressor case includes a first compressor sectionand a second compressor section. A turbine section drives at least oneof the first and second compressor sections, and a gear arrangement isdriven by the turbine section such that the gear arrangement drives thefan section. A plumbing connection area is positioned upstream of thesupport member to be utilized for maintenance and repair.

The various features and advantages of this invention will becomeapparent to those skilled in the art from the following detaileddescription of an embodiment. The drawings that accompany the detaileddescription can be briefly described as follows.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 illustrates a schematic sectional view of an embodiment of a gasturbine engine.

FIG. 2 illustrates a sectional view of a prior art compressor casemounting arrangement. Notably, some aspects are not prior art.

FIG. 3 illustrates a sectional view of an example compressor casemounting arrangement of an embodiment of the current invention.

FIG. 4 illustrates a close up sectional view of the intersection betweenan inlet case and a low pressure compressor case in the embodiment ofFIG. 3 .

FIG. 5 graphically shows a split in the compression ratios between thelow pressure and high pressure compressor sections in a gas turbineengine embodiment.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates an example gas turbine engine 10including (in serial flow communication) a fan section 14, a compressorsection 19 that includes a low pressure (or first) compressor section 18and a high pressure (or second) compressor section 22, a combustor 26,and a turbine section 21 that includes a high pressure (or second)turbine section 30 and a low pressure (or first) turbine section 34. Thegas turbine engine 10 is circumferentially disposed about an enginecenterline X. During operation, air is pulled into the gas turbineengine 10 by the fan section 14, pressurized by the compressors 18, 22mixed with fuel, and burned in the combustor 26. Hot combustion gasesgenerated within the combustor 26 flow through high and low pressureturbines 30, 34, which extract energy from the hot combustion gases. Asused herein, a “high pressure” compressor or turbine experiences ahigher pressure that a corresponding “low pressure” compressor orturbine.

In a two-spool design, the high pressure turbine 30 utilizes theextracted energy from the hot combustion gases to power the highpressure compressor 22 through a high speed shaft 38, and a low pressureturbine 34 utilizes the energy extracted from the hot combustion gasesto power the low pressure compressor 18 and the fan section 14 through alow speed shaft 42. However, the invention is not limited to thetwo-spool gas turbine architecture described and may be used with otherarchitectures such as a single-spool axial design, a three-spool axialdesign and other architectures. That is, there are various types of gasturbine engines, many of which could benefit from the examples disclosedherein, which are not limited to the design shown.

The example gas turbine engine 10 is in the form of a high bypass ratioturbine engine mounted within a nacelle or fan casing 46, whichsurrounds an engine casing 50 housing a core engine 54. A significantamount of air pressurized by the fan section 14 bypasses the core engine54 for the generation of propulsion thrust. The airflow entering the fansection 14 may bypass the core engine 54 via a fan bypass passage 58extending between the fan casing 46 and the engine casing 50 forreceiving and communicating a discharge airflow F1. The high bypass flowarrangement provides a significant amount of thrust for powering anaircraft.

The gas turbine engine 10 may include a geartrain 62 for controlling thespeed of the rotating fan section 14. The geartrain 62 can be any knowngear system, such as a planetary gear system with orbiting planet gears,a planetary system with non-orbiting planet gears or other type of gearsystem. The low speed shaft 42 may drive the geartrain 62. In thedisclosed example, the geartrain 62 has a constant gear ratio. It shouldbe understood, however, that the above parameters are only exemplary ofa contemplated geared gas turbine engine 10. That is, aspects of theinvention are applicable to traditional turbine engines as well as otherengine architectures.

The engine 10 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 10 bypass ratio is greater than about six(6), with an example embodiment being greater than ten (10), the gearedarchitecture 62 is an epicyclic gear train, such as a planetary gearsystem or other gear system, with a gear reduction ratio of greater thanabout 2.3 and the low pressure turbine 34 has a pressure ratio that isgreater than or equal to about 5. In one example, the gearedarchitecture 62 includes a sun gear, a ring gear, and intermediate gearsarranged circumferentially about the sun gear and intermeshing with thesun gear and the ring gear. The intermediate gears are star gearsgrounded against rotation about the axis X. The sun gear is supported bythe low speed shaft 38, and the ring gear is interconnected to the fan14.

In one disclosed embodiment, the engine 10 bypass ratio is greater thanabout ten (10:1), the fan diameter is significantly larger than that ofthe low pressure compressor 18, and the low pressure turbine 34 has apressure ratio that is greater than or equal to about 5:1. Low pressureturbine 34 pressure ratio is pressure measured prior to inlet of lowpressure turbine 34 as related to the pressure at the outlet of the lowpressure turbine 34 prior to an exhaust nozzle. The geared architecture62 may be an epicycle gear train, such as a planetary gear system orother gear system, with a gear reduction ratio of greater than about2.3:1, and more specifically greater than about 2.6:1. It should beunderstood, however, that the above parameters are only exemplary of oneembodiment of a geared architecture engine and that the presentinvention is applicable to other gas turbine engines including directdrive turbofans.

A significant amount of thrust is provided by a bypass flow through thebypass passage 58 due to the high bypass ratio. The fan section 14 ofthe engine 10 is designed for a particular flight condition—typicallycruise at about 0.8 Mach and about 35,000 feet. The flight condition of0.8 Mach and 35,000 ft, with the engine at its best fuelconsumption—also known as “bucket cruise Thrust Specific FuelConsumption (‘TSFC’)”—is the industry standard parameter of lbm of fuelbeing burned divided by lbf of thrust the engine produces at thatminimum point. “Low fan pressure ratio” is the pressure ratio across thefan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The lowfan pressure ratio as disclosed herein according to one non-limitingembodiment is less than about 1.45. “Low corrected fan tip speed” is theactual fan tip speed in ft/sec divided by an industry standardtemperature correction of [((Tambient deg R)/518.7){circumflex over( )}0.5]. The “Low corrected fan tip speed” as disclosed hereinaccording to one non-limiting embodiment is less than about 1150ft/second. The above parameters for the engine 20 are intended to beexemplary.

As shown in FIG. 2 , the example engine casing 50 generally includes atleast an inlet case portion 64, a low pressure compressor case portion66, and an intermediate case portion 76. The inlet case 64 guides air tothe low pressure compressor case 66. The low pressure compressor case 66in an example prior art gas turbine engine 80 supports a plurality ofcompressor stator vanes 68. Notably, the low pressure compressor section18, and the high pressure compressor section 22, and the arrangement ofthe low rotor 70 and high rotor 170, respectively, are not part of theprior art. The low rotor 70 rotates about the central axis X, and, withthe compressor stator vanes 68, help compress air moving through the lowpressure compressor case 66. Downstream of the low pressure compressorthe air passes into the high pressure compressor section 22, and isfurther compressed by its rotor 170. The mounting of the compressor asshown in FIG. 2 is prior art, however, the structure of the low pressurecompressor section 18 and high pressure compressor section 22, and therotors 70 and 170 were not part of the prior art.

A plurality of guide vanes 72 secure the intermediate case 76 to the fancasing 46. Formerly, the guide vanes 72 each included at least a rearattachment 74 and a forward attachment 78. The rear attachment 74connects to an intermediate case 76 while the forward attachment 78connects to the inlet case 64. The lower pressure compressor case 66 wasthus supported through the intermediate case 76 and the inlet case 64.

In the prior art, a plumbing connection area 82 is positioned betweenthe rear attachment 74 and the forward attachment 78. The plumbingconnection area 82 includes connections used for maintenance and repairof the gas turbine engine 80, such as compressed air attachments, oilattachments, etc. The forward attachment 78 extends to the inlet case 64from at least one of the guide vanes 72 and covers portions of theplumbing connection area 82. A fan stream splitter 86, a type of cover,typically attaches to the forward attachment 78 to shield the plumbingconnection area 82.

Referring now to an example of the present invention shown in FIG. 3 ,in the turbine engine 90, the forward attachment 78 attaches to a frontportion of the low pressure compressor case 66. In this example, theforward attachment 78 extends from the guide vane 72 to support the lowpressure compressor case 66. Together, the forward attachment 78 andguide vane 72 act as a support member for the low pressure compressorcase 66. The plumbing connection area 82 (which includes connectionsused for maintenance and repair of the gas turbine engine 90, such ascompressed air attachments, oil attachments, etc) is positioned upstreamof the forward attachment 78 facilitating access to the plumbingconnection area 82. In contrast, the plumbing connection area of priorart embodiments was typically positioned between the rear attachment andthe forward attachment and the forward attachment typically extended tothe inlet case from at least one of the guide vanes, thereby coveringportions of the plumbing connection area, which complicated accessthereto; this complicated structure was further complicated by a fanstream splitter, a type of cover, that typically was attached to theforward attachment to shield the plumbing connection area.

In the embodiment shown in FIG. 3 , an operator may directly access theplumbing connection area 82 after removing the fan stream splitter 86.The plumbing connection area 82 typically provides access to alubrication system 82 a, a compressed air system 82 b, or both. Thelubrication system 82 a and compressed air system 82 b are typically influid communication with the geartrain 62.

Maintenance and repair of the geartrain 62 may require removing thegeartrain 62 from the engine 90. Positioning the plumbing connectionarea 82 ahead of the forward attachment 78 simplifies maintenance andremoval of the geartrain 62 from other portions of the engine 90.Draining oil from the geartrain 62 prior to removal may take placethrough the plumbing connection area 82 for example. The plumbingconnection area 82 is typically removed with the geartrain 62. Thus, thearrangement may permit removing the geartrain 62 on wing or removing theinlet case 64 from the gas turbine engine 90 separately from the lowpressure compressor case 66. This reduces the amount of time needed toprepare an engine for continued revenue service, saving an operator bothtime and money.

Connecting the forward attachment 78 to the low pressure compressor case66 helps maintain the position of the rotor 70 relative to the interiorof the low pressure compressor case 66 during fan rotation, even if thefan section 14 moves. In this example, the intermediate case 76 supportsa rear portion of the low pressure compressor case 66 near a compressedair bleed valve 75.

As shown in FIG. 4 , a seal 88, such as a “W” seal, may restrict fluidmovement between the inlet case 64 and the low pressure compressor case66. In this example, the seal 88 forms the general boundary between theinlet case 64 and the low pressure compressor case 66, while stillallowing some amount of movement between the cases.

FIG. 5 shows a novel worksplit that has been invented to improve thefuel burn efficiency of a geared turbofan architecture with a fan 14connected to the low compressor 18 through a speed reduction device suchas a gearbox 62. Since a gear reduction 62 is incorporated between thefan 14 and the low pressure compressor 18, the speeds of the lowpressure compressor can be increased relative to a traditional two spooldirect drive arrangement. This provides freedom in splitting the amountof compression between the low pressure section 18 and the high pressuresection 22 that can be uniquely exploited to improve fuel burnefficiency on the geared turbofan architecture described in FIGS. 1 and2 . This resulting worksplit is distinctly different from historical twoand three spool direct drive architectures as shown in FIG. 5 .

Notably, while the gear train 62 is shown axially adjacent to the fan14, it could be located far downstream, and even aft of the low turbinesection 34. As is known, the gear illustrated at 62 in FIGS. 2 and 3could result in the fan 14 rotating in the same, or the oppositedirection of the compressor rotors 70 and 170.

It is known in prior art that an overall pressure ratio (when measuredat sea level and at a static, full-rated takeoff power) of at least 35:1is desirable, and that an overall pressure ratio of greater than about40:1 and even about 50:1 is more desirable. That is, after accountingfor the fan 14 pressure rise in front of the low pressure compressor 18,the pressure of the air entering the low compressor section 18 should becompressed as much or over 35 times by the time it reaches the outlet ofthe high compressor section 22. This pressure rise through the low andhigh compressors will be referred to as the gas generator pressureratio.

FIG. 5 shows the way that this high pressure ratio has been achieved inthe two prior art engine types versus the Applicant's engine'sconfiguration.

Area S₁ shows the typical operation of three spool arrangementsdiscussed the Background Section. The pressure ratio of the lowcompressor (i.e., the pressure at the exit of the low pressurecompressor divided by the pressure at the inlet of the low pressurecompressor) is above 8, and up to potentially 15. That is, if a pressureof 1 were to enter the low pressure compressor, it would be compressedbetween 8 to 15 times.

As can be further seen, the high pressure compressor ratio (i.e., thepressure at the exit of the high pressure compressor divided by thepressure at the inlet of the high pressure compressor) in thisarrangement need only compress a very low pressure ratio, and as low as5 to achieve a combined gas generator pressure ratio of above 35. Forexample, if the low pressure compressor ratio is 10 and the highpressure compressor ratio is 3.5, the combined overall pressure ratio(“OPR”) would be (10)(3.5)=35. In addition, the three spool designrequires complex arrangements to support the three concentric spools.

Another prior art arrangement is shown at area S₂. Area S₂ depicts thetypical pressure ratio split in a typical two spool design with a directdrive fan. As can be seen, due to the connection of the fan directly tothe low pressure compressor, there is little freedom in the speed of thelow pressure compressor. Thus, the low pressure compressor can only do asmall amount of the overall compression. As shown, it is typically below4 times. On the other hand, the high pressure compressor must provide anamount of compression typically more than 20 times to reach an OPR of 40(or 50).

The S₂ area results in undesirably high stress on the high pressurecompressor, which, in turn, yields challenges in the mounting of thehigh pressure spool. In other words, the direct drive system thatdefines the S₂ area presents an undesirable amount of stress, and anundesirable amount of engineering required to properly mount the highpressure spool to provide such high pressure ratios.

Applicant's current low compressor/high compressor pressure split isshown at area S₃. The fan is driven at a speed distinct from the lowpressure compressor, and a higher compression ratio can be achieved atthe low pressure compressor section than was the case at area S₂. Thus,as shown, the pressure ratio across the low pressure compressor may bebetween 4 and 8. This allows the amount of compression to be performedby the high pressure compressor to only need to be between 8 times and15 times.

The area S₃ is an enabling design feature that allows the gearedturbofan architecture shown in FIGS. 1 and 2 to achieve a very high gasgenerator OPR while avoiding the complexities of historical three spooland two spool direct drive architectures. The area S₃ is an improvementover both areas S₁ and S₂. As an example, a 3-4% fuel efficiency isachieved at area S₃ compared to area S₁. A fuel savings of 4-5% isachieved at area S₃, compared to area S₂.

In fact, in comparison to a gas turbine engine provided with a geardrive, but operating in the pressure ratios of area S₂, there is still a2% fuel burn savings at the S₃ area.

As such, the area S₃ reduces fuel burn, and provides engineeringsimplicity by more favorably distributing work between the hotter highpressure spools and colder low pressure spools.

Stated another way, the present invention provides a combination of alow pressure compressor and a high pressure compressor which togetherprovides an OPR of greater than about 35 and, in some embodimentsgreater than about 40, in some embodiments greater than about 50, and insome embodiments up to about 70. This high OPR is accomplished by abeneficial combination of a pressure ratio across the low pressurecompressor of between about 4 and about 8 coupled with an additionalpressure ratio across the high pressure ratio compressor of betweenabout 8 and about 15.

Improved fuel consumption can be further achieved wherein the fan may below pressure, and have a pressure ratio less than or equal to about1.50, more specifically less than or equal to about 1.45, and even morespecifically less than or equal to about 1.35. A bypass ratio, definedas the volume of air passing into bypass passage 58 compared to thevolume of air in the core air flow is greater than or equal to about 8at cruise power. The low pressure compressor may have a pressure ratioless than or equal to 8, more narrowly between 3 to 8, and even morenarrowly 4 to 6, and be powered by a 4 or 5-stage low pressure turbine.In some embodiments, the first or low pressure compressor may have apressure ratio greater than or equal to 7. The second or high compressorrotor may have a nominal pressure ratio greater than or equal to 7, morenarrowly between 7 to 15, and even more narrowly 8 to 10, and may bepowered by a 2-stage high pressure turbine. A gas turbine engineoperating with these operational parameters provides benefits comparedto the prior art.

Although an embodiment of this invention has been disclosed, a worker ofordinary skill in this art would recognize that certain modificationswould come within the scope of this invention. For that reason, thefollowing claims should be studied to determine the true scope andcontent of this invention.

We claim:
 1. A turbofan gas turbine engine comprising: a fan sectionincluding a fan having at least one fan blade, with a fan pressure ratioof less than 1.45, the fan pressure ratio measured across the fan bladealone, and a fan casing surrounding the fan to define a bypass passage;a core engine including a compressor section and a turbine section, anda bypass ratio defined as the volume of air passing into the bypasspassage compared to the volume of air passing into the core engine,wherein the bypass ratio is greater than 10; a gear arrangement thatdrives the fan section, the gear arrangement including an epicyclic geartrain having a sun gear, a ring gear, and a plurality of intermediategears arranged circumferentially about the sun gear and intermeshingwith the sun gear and the ring gear, and the gear arrangement defining agear reduction ratio of greater than 2.3:1; a lubrication system and acompressed air system in fluid communication with the gear arrangement;wherein the compressor section includes a three-stage low pressurecompressor section and a high pressure compressor section; wherein theturbine section drives the gear arrangement, the turbine sectionincludes a low pressure turbine and a two-stage high pressure turbine,the high pressure turbine drives the high pressure compressor section,the low pressure turbine includes an inlet, an outlet, and a lowpressure turbine pressure ratio greater than 5:1, wherein the lowpressure turbine pressure ratio is a ratio of a pressure measured priorto the inlet as related to a pressure at the outlet prior to any exhaustnozzle, and the high pressure compressor section has a greater number ofstages than a total number of stages of the turbine section; wherein anoverall pressure ratio is: provided by the combination of a pressureratio across the low pressure compressor section and a pressure ratioacross the high pressure compressor section; and greater than 40;wherein the pressure ratio across the high pressure compressor sectionis between 7 and 15; and wherein the pressure ratio across the lowpressure compressor section is between 4 and
 8. 2. The turbofan gasturbine engine of claim 1, wherein the overall pressure ratio is greaterthan
 50. 3. The turbofan gas turbine engine of claim 2, wherein the lowpressure turbine drives the low pressure compressor section and the geararrangement, and the gear arrangement is incorporated between the fanand the low pressure compressor section.
 4. The turbofan gas turbineengine of claim 3, wherein the fan has a low corrected fan tip speed ofless than 1150 feet/second.
 5. The turbofan gas turbine engine of claim4, wherein the overall pressure ratio is no greater than
 70. 6. Theturbofan gas turbine engine of claim 5, wherein the gear train has aconstant gear ratio.
 7. The turbofan gas turbine engine of claim 6,wherein the pressure ratio across the low pressure compressor section isgreater than or equal to
 7. 8. The turbofan gas turbine engine of claim6, wherein the pressure ratio across the high pressure compressorsection is above
 10. 9. The turbofan gas turbine engine of claim 6,wherein the fan pressure ratio is less than or equal to 1.35.
 10. Theturbofan gas turbine engine of claim 5, further comprising a centralaxis, wherein the intermediate gears are star gears grounded againstrotation about the central axis, and the ring gear is interconnected tothe fan.
 11. The turbofan gas turbine engine of claim 10, wherein thepressure ratio across the low pressure compressor section is between 4and 6, and the pressure ratio across the high pressure compressorsection is between 8 and
 10. 12. The turbofan gas turbine engine ofclaim 11, wherein the low pressure turbine is a four-stage or five-stageturbine, and the high pressure compressor section is a nine-stagecompressor.
 13. The turbofan gas turbine engine of claim 5, wherein thegear arrangement is a planetary gear system, and the intermediate gearsare orbiting planet gears.
 14. The turbofan gas turbine engine of claim13, wherein the pressure ratio across the low pressure compressorsection is between 4 and 6, and the pressure ratio across the highpressure compressor section is between 8 and
 10. 15. The turbofan gasturbine engine of claim 14, wherein the low pressure turbine is afour-stage or five-stage turbine, and the high pressure compressorsection is a nine-stage compressor.
 16. A turbofan gas turbine enginecomprising: a fan section including a fan having at least one fan blade,with a fan pressure ratio of less than 1.45, the fan pressure ratiomeasured across the fan blade alone, and a fan casing surrounding thefan to define a bypass passage; a core engine including a compressorsection and a turbine section, and a bypass ratio defined as the volumeof air passing into the bypass passage compared to the volume of airpassing into the core engine, wherein the bypass ratio is greater than10; a gear arrangement that drives the fan section, the gear arrangementincluding an epicyclic gear train having a sun gear, a ring gear, and aplurality of intermediate gears arranged circumferentially about the sungear and intermeshing with the sun gear and the ring gear, and the geararrangement defining a gear reduction ratio of greater than 2.3:1; alubrication system and a compressed air system in fluid communicationwith the gear arrangement; wherein the compressor section includes a lowpressure compressor section and a nine-stage high pressure compressorsection; wherein the turbine section drives the gear arrangement, theturbine section includes a four-stage low pressure turbine and atwo-stage high pressure turbine, the high pressure turbine drives thehigh pressure compressor section, the low pressure turbine includes aninlet, an outlet, and a low pressure turbine pressure ratio greater than5:1, wherein the low pressure turbine pressure ratio is a ratio of apressure measured prior to the inlet as related to a pressure at theoutlet prior to any exhaust nozzle, and the low pressure compressorsection has a greater number of stages than the high pressure turbine;wherein an overall pressure ratio is: provided by the combination of apressure ratio across the low pressure compressor section and a pressureratio across the high pressure compressor section; and greater than 40;wherein the pressure ratio across the high pressure compressor sectionis between 7 and 15; and wherein the pressure ratio across the lowpressure compressor section is between 4 and
 8. 17. The turbofan gasturbine engine of claim 16, wherein the overall pressure ratio isgreater than
 50. 18. The turbofan gas turbine engine of claim 17,wherein the low pressure turbine drives the low pressure compressorsection and the gear arrangement, and the gear arrangement isincorporated between the fan and the low pressure compressor section.19. The turbofan gas turbine engine of claim 18, wherein the fan has alow corrected fan tip speed of less than 1150 feet/second.
 20. Theturbofan gas turbine engine of claim 19, wherein the overall pressureratio is no greater than
 70. 21. The turbofan gas turbine engine ofclaim 20, wherein the gear train has a constant gear ratio.
 22. Theturbofan gas turbine engine of claim 21, wherein the pressure ratioacross the low pressure compressor section is greater than or equal to7.
 23. The turbofan gas turbine engine of claim 21, wherein the pressureratio across the high pressure compressor section is above
 10. 24. Theturbofan gas turbine engine of claim 21, wherein the fan pressure ratiois less than or equal to 1.35.
 25. The turbofan gas turbine engine ofclaim 20, further comprising a central axis, wherein the intermediategears are star gears grounded against rotation about the central axis,and the ring gear is interconnected to the fan.
 26. The turbofan gasturbine engine of claim 25, wherein the pressure ratio across the lowpressure compressor section is greater than or equal to
 7. 27. Theturbofan gas turbine engine of claim 25, wherein the pressure ratioacross the low pressure compressor section is between 4 and 6, and thepressure ratio across the high pressure compressor section is between 8and
 10. 28. The turbofan gas turbine engine of claim 20, wherein thegear arrangement is a planetary gear system, and the intermediate gearsare orbiting planet gears.
 29. The turbofan gas turbine engine of claim28, wherein the pressure ratio across the low pressure compressorsection is greater than or equal to
 7. 30. The turbofan gas turbineengine of claim 28, wherein the pressure ratio across the low pressurecompressor section is between 4 and 6, and the pressure ratio across thehigh pressure compressor section is between 8 and 10.